1. Technical Field
The present invention relates to a rocket engine.
2. Background of the Invention
During operation, a rocket combustion chamber or a rocket engine nozzle is subjected to very high stresses, for example in the form of a very high temperature on its inside (on the order of magnitude of 980° F.) and a very low temperature on its outside (on the order of magnitude of −370° F.). As a result of this high thermal load, stringent requirements are placed upon the choice of material, design and manufacture of the outlet nozzle. At a minimum, the need for effective cooling of the combustion chamber or the outlet nozzle must be considered.
It is a problem to construct cooled wall structures that are capable of containing and accelerating the hot exhaust gas, and also be reliable through a large number of operational cycles. Known designs do not have a sufficiently long service life required to withstand a large number of operational cycles. These known systems generate large thermal stresses, including large pressure drops, or present difficulties when needing repair.
When applying the expander engine cycle, there is a secondary problem. The expander engine cycle uses the cooling medium to drive the turbines in the fuel and oxidator turbo pumps; that is, energy from the expansion of the heated cooling medium is used for driving the turbines. The efficiency of the rocket engine is a function of the combustion pressure. To reach high pressure experience in the expander cycle, efficient heat transfer from the exhaust gas to the cooling medium is required. Increase in the heat load in the combustion chamber due to surface roughness or fins may impair the service life of the engine since the intensity of the heat load is very high in the combustion chamber. Still further, a longer combustion chamber increases the length of the engine and the rocket. A commensurate increase in the size of the nozzle gives rise to larger nozzles and longer rocket structures, each of which increases the weight of the vehicle.
There are several different known methods for manufacturing a rocket nozzle with cooling channels. According to one of these methods, the nozzle has a brazed tube wall. The tubes have a varying cross-sectional width to provide the contour of the nozzle when assembled. The variation in cross section is given by variation of the circumference and by variation of the form of the cross section. The brazed joints restrict the deformation of tubes in the thermal expansion and pressure cycle. The stresses in the tubes are increased in the arc of the joints. The joints themselves are weak points that may break and are difficult to repair. The brazed tube wall provides a larger “wet” contact surface for the rocket flame than a sandwich wall or a constant tube section wall. However, even larger wet surfaces are desirable.
According to another known method, a sandwich wall is made by milling channels in sheet metal and joining a thinner sheet metal to seal the channels. The inner and outer walls are continuous shells. In the thermal cycle, the walls are in compressive and tension strain. This type of wall structure is not well suited to sustain the tension loads normal in the service life of a rocket nozzle. The sandwich wall features no increase in surface area to enhance heat transfer.
According to still another known method, the nozzle wall is manufactured with constant section tubes. The tubes are helically wound and welded together to form the nozzle contour. The increase in surface area is small. The tubes have an angle relative to exhaust gas flowing through the nozzle. This helps to increase the heat transfer. However, at the same time the exhaust flow is rotated and a reactive roll momentum influences the flight of the rocket. The constant section tubes result in a large pressure drop that is not favorable for convectively cooled engines. The large pressure drop is negative for the expander cycle type engine.